1. Field of the Invention
The present invention relates generally to the field of composite structures and, more particularly but not by way of limitation, to composite structures having skins separated and stiffened by hollow hats, each of which incorporates an integral co-cured fly away hollow mandrel used in laying up and curing the structure.
2. Prior Art
There is a growing trend in the aerospace industry to expand the use of advanced composite materials for a diverse array of structural and dynamic aerostructural applications because of the strength-to-weight advantage provided by composite materials. One particular application for the use of such advanced composite materials, such as graphite or an aromatic polyamide fiber of high tensile strength that are embedded in a resinous matrix, e.g., an epoxy, is for airfoil structures that are composed of skins separated and stiffened by a honeycomb core material. In the instance of an aerospace article such as a fan cowl, one or more stiffening members are affixed to the outer skin and covered with an inner skin for efficiently transmitting and/or reacting axial and/or bending loads to which the fan cowl is subjected.
There are two techniques currently employed for bonding through autoclave processing a composite stiffening member in combination with composite face layers: (1) the secondary bonding method, and (2) the co-cured bonding method. Both methods are disadvantageous in requiring costly non-reusable tooling and/or costly and tedious manufacturing steps.
A typical composite sandwich panel intended for use as an aerostructure part is normally fabricated using two autoclave-cured inner and outer composite skins that are formed by using a curing cycle with heat, pressure, and a unique tool for each skin. A sandwich panel is then made up using a composite bond jig, tool or fixture with the pre-cured face skin laid-up on the bond jig tool followed by a ply of film adhesive. A honeycomb aluminum or non-metallic core of a given thickness is placed on the face skin, another ply of film adhesive is applied, and finally the previously pre-cured inner skin is placed on the adhesive film. The bond jig that is used to fabricate the sandwich panel is usually the same tool that was used to create the outer composite skin. A plurality of closure plies of uncured composite material are laid up. Next, the assembled sandwich panel is cured during its final assembly stage. The entire sandwich panel is then vacuum bagged to the composite bond jig and again cured in an autoclave under high pressure and heat.
Thus, at least three very expensive and labor intensive fabrication and cure cycles have gone into the production of the exceptionally strong and lightweight composite honeycomb core sandwich panel. At least two different and expensive tools are needed in this process. Manufacturing flow time is very long, energy use is high, and the manufacturing floor space required is considerable.
The second method referred to above, the co-curing method, involves curing the composite inner and outer skins that are laid up with a layer of adhesive film and honeycomb core in one cure cycle in the autoclave. A co-cured panel is desirable in that it is less expensive to fabricatexe2x80x94only one bond jig tool is required, only one cure cycle is needed, the method is less labor-intensive, less floor space is required, and a much shorter manufacturing flow time is achieved. However, co-curing an aerostructure panel has never achieved wide-spread acceptance because of a large loss of panel strength and integrity, which is due to the lack of compaction of the composite plies placed over and under the honeycomb core. The composite plies dimple into the center of each core cell with nothing but the cell walls to compact the composite skins. The only way to overcome this xe2x80x9cknockdownxe2x80x9d characteristic is to add extra plies, which creates both unwanted weight and added cost. Thus, because of these constraints co-cured aerostructure panels are not widely manufactured in the aerospace industry.
There are other particular problems when a honeycomb core element is used to provide a stiffening element for an aerospace component. As Hartz et al. have described in U.S. Pat. No. 5,604,010 concerning a xe2x80x9cComposite Honeycomb Sandwich Structure,xe2x80x9d with a high flow resin system large amounts of resin can flow into the core during the autoclave processing cycle. Such flow robs resin from the laminate, introduces a weight penalty in the panel to achieve the desired performance, and forces over-design of the skin plies to account for the resin losses to the honeycomb core. To achieve the designed performance and the corresponding laminate thickness, additional plies are necessary with resulting cost and weight penalties. Because the weight penalty is severe in terms of the impact on vehicle performance and costly in modern aircraft and because the resin flow is a relatively unpredictable and uncontrolled process, aerospace design and manufacture dictates that flow into the core be eliminated or significantly reduced. In addition to the weight penalty from resin flow to the core, it has been learned that micro-cracking that originated in the migrated resin can propagate to the bond line and degrade mechanical performance. Such micro-cracking potential has a catastrophic threat to the integrity of the panel and dictates that flow be eliminated or at least controlled.
Unfortunately, the use of a honeycomb core as a stiffener for elements in a aerostructure component, such as a structural panel, has other deleterious effects. Two of the greatest drawbacks to an aluminum core are its inherent significant cost and susceptibility to corrosion. To minimize galvanic corrosion of the core caused by contact with the face skins, isolating sheets are interposed between the aluminum core and the face skins. Also, the aluminum core is expensive and also must be machined to a desired shape in a costly process. The honeycomb core may also be subject to crush during manufacture, which imposes a limit on the pressures that may be used in autoclave processing. Thus, the processing of an aerospace advanced composite article is limited to an autoclave pressure of not greater than 45 psi, rather than a higher pressure that would increase the strength of the resultant advanced composite article. Also, the honeycomb core, if damaged in use, has a spring-back property, which makes the detection of such damage more difficult.
In providing reinforcing mandrels for stiffener elements, such as hat sections, for aerospace advanced composite structural panels, it is also known to provide a composite stiffening member in the form of a polyamide foam mandrel fabricated by machining a core mandrel to a desired shape. Obviously, the machining of the core mandrel is expensive and time consuming and further introduces the problem of properly bonding the core mandrel to the inner and outer skins.
Therefore, a great need has arisen for a practical method of readily producing stiffened, fiber-reinforced composite structures useful in the construction of integrally stiffened components for aerospace applications, which are cost and labor efficient and which save time in the fabrication process.
Accordingly, it is an object of the present invention to provide a method for fabricating aerostructure advanced composite articles that eliminates a honeycomb core as a spacing and stiffening element, provides a lighter weight assembly, and is easier to repair. Another object of the present invention is to reduce the lay-up cost of known advanced composite co-cure assemblies by at least 15% and to increase assembly strength over previously known co-cure assembly methods by being able to utilize high pressures in autoclave processing. Yet another object of the present invention is to improve the quality of co-cured advance composite assemblies and thereby increase customer satisfaction. A further object of the present invention is to provide a process that provides an assembly that can be manufactured in one manufacturing cell from raw material to final product. Yet another object of the present invention is to reduce the cost of post-bond and final assembly work for the final co-cured assembly, which assembly will readily indicate damage to an improved stiffening element.
The foregoing discussion covers some of the more significant objects of the invention. Those objects should be construed to be merely illustrative of some of the more prominent features and applications of the present invention. Many other beneficial results can be attained by applying the invention in a different manner or by modifying the invention within the scope of the disclosure. Accordingly, other objects and a fuller understanding of the invention may be had by referring to the summary of the invention and the detailed description of the preferred embodiments in addition to the scope of the invention defined by the claims taken in conjunction with the accompanying drawings.
The foregoing objects are attained, in accordance with one aspect of the present invention, by a composite structure having first and second composite skin layers, each of a resin-impregnated fiber material, and two or more elongated stiffener/spacer composite members, each of a resin-impregnated fiber material and interposed between the skin layers. The stiffener/spacer composite members are arranged generally longitudinally of the skin layers in spaced-apart relation laterally. Each stiffener/spacer composite member has side walls extending between and united with the skin layers by base walls that are bonded to the inner faces of the skin layers. An elongated hollow mandrel of a stiffened fabric is received between and in engagement with the side walls of each the of the stiffener/spacer composite members, the mandrel serving as a form for shaping the stiffener/spacer composite member. The composite layers are co-cured under a predetermined pressure and a predetermined temperature to render the structure unitary.
As will be more readily apparent from the description below, the composite structure of the present invention provides very strong sandwich aerostructures having co-cured skins and either hat-shaped or box-shaped tubular stiffener/spacers connecting the skins. The presence of the hollow mandrels, which enable the stiffener/spacers to be laid up from uncured resin-impregnated fiber material and assembled with the skin layers of likewise uncured composite skin layers, avoids the use of honeycomb cores and other costly core materials in the aerostructure and in addition to reducing costs allows significant weight savings. The hollow mandrels also leave open spaces within the stiffener/spacer members, which facilitate installing connectors, fittings, hydraulic lines and electric wiring within the structures. The side walls and the base walls of the stiffener/spacer members stiffen the skins and provide considerable versatility in the structural and aerodynamic design of the composite structure.
The stiffener/spacer composite members may be generally hat-shaped in cross section with a generally U-shaped body portion, the legs of which form the walls, and a side flange portion forming one of said base walls extending from the end of each of the side walls of the U-shaped body portion. Alternatively, the stiffener/spacer composite members may be tubular with the side walls joined by at least one of the base walls. The tubular stiffener/spacer composite members may have a generally U-shaped body portion and inturned flange portions forming one of the base walls. The inner edges of the flange portions may meet at a butt joint or form a gap. Tubular stiffener/spacer composite members may also be formed by helical plies of composite material wrapped around the mandrel. Each hollow mandrel is, preferably, formed of a unitary stiffened carbon fabric tape.
Variations of the placement and geometry of the hollow mandrels permit the composite structure to be of various shapes and of various structural forms. One or more of the stiffener/spacer composite members may be of uniform cross section along at least a portion or the entirety of its length. One or more of the stiffener/spacer composite members may be of non-uniform cross section along at least a portion or the entirety of its length.
At least a portion of one or more of the stiffener/spacer composite members may have a longitudinal axis that is curved. The width of one or both base walls of one or more of the stiffener/spacer composite members may vary along at least a portion or along the entirety of its length. Similarly, the heights of the side walls of the stiffener/spacer composite members may be the same or they may be different along at least portions or along the entireties of their lengths. The stiffener/spacer members may be parallel or they may converge longitudinally along at least portions or along the entireties of their lengths.
The foregoing variations in the cross-sectional and longitudinal geometries and placements of the stiffener/spacer members permit the skin layers to be flat or partly or fully curved in either the lateral or the longitudinal direction or both and to be spaced apart at constant or varying distances.
In a particularly useful configuration for aerostructures forming airfoils, the top and bottom skin layers may meet and be in engagement at a longitudinal edge juncture. Advantageously, an elongated composite member of wedge shape in cross section and of a resin-impregnated fiber material is received between and engaged with portions of the top and bottom skin layers adjacent the edge juncture.
According to another aspect of the present invention, a method is provided for fabricating a composite structure from fiber-reinforced composite materials that includes the steps of assembling on a forming surface of a first jig a first uncured composite skin layer of a resin-impregnated fiber material, assembling on a forming surface of a second jig a second uncured composite skin layer of a resin-impregnated fiber material, and providing first and second elongated hollow mandrels, each hollow mandrel being of a stiffened fabric and having a predetermined shape in cross section. A stiffener/spacer composite member of an uncured resin-impregnated fiber material is assembled over each mandrel to form side walls and base walls. Each assembly of a stiffener/spacer composite member over a mandrel is combined with one of the skin members such that one of the base walls is in engagement with that skin member. The jigs are then juxtaposed in opposed relation to form a sandwich of the skin layers and stiffener/spacer composite members between the jigs. A hollow vacuum tube bag is placed into each mandrel and into the space between the side walls of the stiffener/spacer members assembled over the hollow mandrels. A vacuum bag is placed over the first jig and the sandwich. The tube bags are sealed to the vacuum bag, the second jig and each other in a manner such as to permit the hollow tube bags to communicate with the exterior of the vacuum bag. The vacuum bag is evacuated and the sandwich is subjected to a cycle of predetermined pressure and temperature to cure the resin-impregnated fiber materials of the composite skin layers and the stiffener/spacer composite members.
The method of the invention has the distinct advantage of allowing the structures forming the sandwich to be laid up in the uncured states and co-curedxe2x80x94separate curing of the skin layers, followed by combining the skin layers with a core structure and a second curing step, is avoided, with significant time, cost and energy savings. The mandrels used as forms for laying up the stiffener/spacer members can be provided in a wide variety of shapes and allow the fabrication of composite sandwich structures with the features described above.
In preferred embodiments of the method, each hollow mandrel is formed of a unitary stiffened carbon fabric tape that is spirally arranged to extend longitudinally and to provide a predetermined cross section.
It is highly advantageous to cure the sandwich using vacuum-bagging technology, as described generally above, which allows curing under pressures above about 45 psi, and preferably of about 70 psi, to compress the layers during curing for enhanced strength of the structure. When the vacuum-bagged jigs and sandwich of composite skins and stiffener/spacer composite members are placed in an autoclave for curing, the elevated pressure is exerted externally on the main vacuum bag on one side of the sandwich and on the jig on the other side of the sandwich and internally within each of the vacuum tube bags. The pressure acting internally on the top and bottom walls of the tube bags compresses the skin layers and the base walls of the stiffener/spacer composite members against the inner forming surfaces of the jigs. The pressure acting internally on the Bide walls of the vacuum tube bags compresses the side walls of the stiffener/spacer members. One of the beauties of the vacuum-bagging technique as applied in the present invention is that the opposite layers of the composite material are subjected to equal pressure, so there is no tendency to deform the layers. Stability of the side walls of the stiffener/spacer members, which are not backed up by conventional rigid xe2x80x9cformxe2x80x9d elements, is aided during autoclaving by the hollow mandrels.
Following the cure cycle the jigs and co-cured sandwich are removed from the autoclave, and the main vacuum bag and the vacuum tube bags are removed. The structure may then be trimmed as desired.
As mentioned above, the stiffener/spacer members may be hat-shaped in cross-section. In that case, a mandrel is first secured to one of the skin layers by an uncured adhesive, which upon curing bonds the mandrels to the portions of the skins to which they are adhered. The stiffener/spacer composite layer is then assembled over the mandrel. Tubular stiffener/spacer composite members may be laid up on the mandrels and then assembled as a unit to one of the composite skin layers.
The foregoing description has outlined rather broadly some features and advantages of the present invention. The detailed description of embodiments of the invention that follows will enable the present invention to be better understood and the present contribution to the art to be more fully appreciated. Those skilled in the art will recognize that the embodiments may be readily utilized as a basis for modifying or designing other structures and methods for carrying out the purposes of the present invention. All such structures and methods are intended to be included within the spirit and scope of the invention as set forth in the appended claims.